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How do supersonic warplanes deal with birds?
The Ruppell's Vulture can reach 11 km and can weigh 9kg. The Kori bustard can weigh 18kg. Many warplanes must break the sound barrier within reach of birds, especially on terrain hugging or sea skimming missions. How do they deal with the possibility of a bird causing structural damage or smashing the windshield?

Mainly, it's in the category of "accepted risk". Warplanes accept lots of risks, like flying around in places where they might get shot at, and carrying large amounts of high explosives, that less risk-averse missions, like commercial airlines, would never accept. Flying low & fast is judged to contribute more to survival in a war-zone than flying placidly along at more bird-free & fuel-efficient altitudes and speeds, even given the hazards of hitting the birds.
Thick, heavy, carefully designed & tested windscreens and canopies serve to lessen the risks, as does having multiple engines and the option to trade airspeed for altitude to get away from the ground rapidly after taking damage from a bird strike, and there are ejection seats if the damage is too serious (but not immediately fatal, obviously).
But, after what steps that can be taken, have been taken, risk remains. And that risk is accepted as part of the military mission.

Leading edges of wings on fighters are pretty sharp and rigid, so they take bird strikes in stride or at least with only minimal deformation, and windshields are designed to withstand projectiles, so same thing there.
The main risk of birds for fighters is ingestion into the engines, and this is a significant risk on a low level strike mission especially in a single engine fighter. Ingest a bird, and unless the design has a specific particle separation feature in its intake ducting that allows a bird to avoid the compressor/fan inlet, or an engine fan with blades that can tolerate a bird carcass strike, you will be ejecting.

An early astronaut Theodore Freeman died in just such a T-38 /goose strike.

fighter jet's rarely go supersonic except in combat conditions or immediate intercept orders. Most birds never fly higher than 500-1000 feet anyway, way below the "Hard deck" operating safety of most fighters. The real risk bird ingestion is at the runway/flight deck. It's simply aspect of calculated risk

re " Most birds never fly higher than 500-1000 feet anyway" -- wrong.  Google diurnal soaring birds, nocturnal avian migration, etc.  If you had phrased it "the majority of bird flight is below 1000'" you might be able to make a case for that but to say "most birds never..." is obviously completely wrong.  But you miss the point of the original question in another way too, as it specifies "especially on terrain hugging or sea skimming missions"--


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How far from the Concorde's ground track was the sonic boom tolerable?
How far away from the ground track did it drop to, say, the level.of heavy traffic? Or something else that people do not mind hearing?

Does this answer your question? Can a sonic boom produced at 60,000' be heard on the ground?

@AadirajAnil no, it's about altitude, not ground track

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How do canards reduce the X-59's sonic boom?
X-59

As of 2017, the ground noise was expected to be around 60 dB(A), about 1/1000 as loud as current supersonic aircraft. This was to be achieved by using a long, narrow airframe and canards to keep the shock waves from coalescing.

I fail to see the relevance of the canards. How are they supposed to reduce the boom?

One of the important factors for sonic booms is the longitudinal distribution of lift along the aircraft.  The canard helps bring some of that lift very far forward.
Notice also that the aircraft has two horizontal tails – the large one in plane with the wing and another smaller one at the top of the vertical tail.  These work to distribute the lift – and thereby the pressure signature – in a tailored way.

I assume the oddly flat nose is for a similar purpose?

The X59 has not flown yet, so there's no confirmed guarantee it will successfully yield a quieter plane. Despite its talk about “quiet” SSTs, Boom is lobbying and affecting legislation that would exempt these and other jets from new stringent noise standards.

The question is specifically about the canards and how they would interact with shockwave intensity.

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How much of the Virgin GlobalFlyer's L/D is due to laminar flow?
Despite travelling much faster than the Voyager, it has a significantly higher L/D. The latter seems to have two features that disrupt laminate flow, that are absent in the former. Namely the canards and the fire propellers.
Anyway, how much of the 37 L/D was due to laminar flow? How much greater would total drag have been without it?


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How can tailplane force imbalance twist a tailcone off?
http://www.rvs.uni-bielefeld.de/publications/Incidents/DOCS/ComAndRep/LaudaAir/LaudaRPT.html#2.2.3

2.2.2 In-Flight Breakup Sequence

The analysis of the major structural damage showed that the failures were probably the result of buffeting, maneuvering overload, and excessive speed. Parts of the airplane that separated from buffeting overload appear to be pieces of the rudder and the left elevator. This was followed by the down-and- aft separation

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of most of the right horizontal stabilizer from maneuvering overloads, as the crew attempted to control the airplane and arrest the high-speed descent. No evidence of impacts were observed on the leading edges of the horizontal and vertical stabilizers indicating that no airframe structural failure occurred prior to horizontal stabilizer separation. It is thought that the download still present on the left stabilizer and the imbalance in the empennage from the loss of the right stabilizer introduced counterclockwise (aft looking forward orientation) torsional overload into the tail, as evidenced by wrinkles that remained visible in the stabilizer center section rear spar. The separation of the vertical and left horizontal stabilizers then occurred, although the evidence was inconclusive as to whether the vertical stabilizer separated prior to or because of the separation of the left stabilizer and center section. (The damage indicated that the vertical stabilizer and the attached upper portion of four fuselage frames departed to the left and that separation of the vertical fin-tip and the dual-sided stringer buckling in the area of the fin-tip failure occurred from bending in both directions prior to the separation of the vertical stabilizer from the fuselage).

I don't understand. The tailcone is already designed to take torsion from the much larger tailfin. And we know that A587's tailfin snapped off rather than twisting the entire tailcone off. So how can a tailplane cause so much stress rather than snapping off?

In most of these airplanes the vertical fin is attached to lugs along the upper skin and frame structure and to the bulkheads at the front and rear spars of the fin. Overloads will tend to break the structure at the interface of the fin and fuselage, the weakest point when subjected to bending loads.
The horizontal tail is quite different. It's normally a one-piece torsion box very similar to the main wing, whose primary structural connection is a massive hinge joint within the tail cone structure at the rear, and a jack screw attachment at the front. The result is torsional overloads of the horizontal tail will tend to brake apart the structure inside the tail cone itself rather than the horizontal stabilizer box or its hinge lugs, whereas the fin tends to want to break where it's attached at the skin/fin interface.


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Solar Impulse lift to drag ratio
What are the L/D ratios of the two Solar Impulse airplanes? I haven't been able to find them.

Solar Impulse1 had 4 × 7.5 kW (10 hp) electric motors and a MTOW of around 4400 pounds, and was able to cruise at 43 mph, with 11 foot propellers.
Solar Impulse 2 had 4 × 13 kW (17.4 hp) each at 5100 lbs, and was able to cruise at 56 mph, with 13 foot propellers.
According to this article, the lift to drag ratio was around 37:1.
Rough calculation of 4 × 17.5 hp cruising at half throttle generating 4 pounds of thrust per hp

5100 lbs/(4×17.5 hp×4 lbs/hp×0.5) = 36.4

This is quite good for powered aircraft$^1$.  Airliners range in the high teens, with the U-2 at 25.6.
$^1$  unless they can retract their power unit into the fuselage


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How close is this Su-27 to the aircraft recording the pass?
There is footage of an intercept of a US Nay recon aircraft being intercepted by a Russian Su-27:

(still from the BBC video article, who look like they got it from the US Navy directly)
which was reported on the BBC, CNN and elswehere.
I am not familiar with the proportions of the intercepting plane, but it looks like it passed very close to the recording aircraft!
Is there a good way to estimate how close it was, to within a rough ballpark? Tens of meters, a hundred meters, three hundred meters?
The CNN article reports that

Officials so far, have not been able to estimate how close the Russian aircraft came to the US plane, but described the flight behavior of the Russians as the key factor in making the determination the encounter was unsafe.

but

The US crew reported turbulence following that initial interaction in which the direct pass occurred. [During the second pass the Su-27] applied its afterburner while conducting a banking maneuver, which is believed to have caused a vibration that was experienced by the American crew.

On a related note, there was also a report of an Su-27 coming within 5 feet (!) of another aircraft earlier in the year; so it seems that this is not a one-off occurrence.

First question here, so I am open to suggestions / improvements :)

Just for comparison, the SU-27 is 72 feet long, so I'd guess that the distance in the picture you posted is 100-150 feet or more.

It looks like it's maybe 50-75 ft away from the camera, but you would have to know whether the camera's focal length is neutral, magnifying or wide angle, so it's really a wild guess.  There's not enough detail on the host aircraft to be able to scale it that way.

This seems like a decent ballpark estimate - I did wonder if having some of the recording aircraft in the video would help with the estimation ("If I know the wing is X feet long and I can see Y % of it taking up Z % of the frame, the amount of frame the Su-27 takes up is..." type approximation; but if there's not enough detail that's fair enough. All that being said, it probably works out about as well as someone with experience eyeballing it anyway! :)


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What is the supersonic L/D of subsonic jetliner
Suppose you somehow managed to accelerate a typical jetliner oastbthe sound barrier and to whatever speed you wanted, climbing to keep lift constant. Is there any data on his the L/D evolves with Mach number in this scenario?


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Why is hybrid laminar flow control preferred for swept wings over other options?
As I understand it, a (the?) major form of laminar-to-turbulent transition on the swept wings of large jets is the crossflow caused by the sweep. One way to control this is laminar flow control by boundary layer suction.
Now, I would expect crossflow to exist over all of the wing chord. Yet, hybrid laminar flow control, the solution I see most commonly proposed for the swept wing, only has suction on the leading edge, followed by natural laminar flow behind. Supposedly this gives simplicity yet most of the benefits of full-chord laminar flow control.
Yet we know that the favorable pressure gradient introduced by natural laminar flow is destabilizing to laminar flow on wings! And that crossflow still exists at the back. So, what makes this only partial suction a good compromise? Is it because the crossflow issue is bigger at the leading edge than elsewhere?


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How high does the "bug layer" extend?
As far as I know, a big obstacle to attaining laminar flow in aircraft today is the so-called "bug layer", presumably a lower layer of the atmosphere where bugs are encountered. So, my question is simply this: how high up do insects stop becoming a serious issue for laminar flow?

Depends on the weather. In a high pressure area air sinks continuously down and so do insects in it. Conversely, low pressure means rising air and carries the insects with it (a bit like gliders).

In 1961 study a plane flew 116,684 miles sampling the air, catching whatever was up there, and managed to capture a single termite at 19,000 feet. This is the highest ever observed.
You just never get above the "bug layer" in a smaller aircraft (Cessna 172 limit is 15,000 feet). Likely this is not a proper naming.

I agree with your suspicion on the naming. +1.

I don't have any scientific data, but in my flying experience, including some time back in the 70s in a Breezy, where, as on a motorcycle, you have to be careful to keep your mouth closed, you don't encounter that many bugs above 1000 ft agl (they are there, but few and far between relatively).  Descending in the Breezy in a warm summer late afternoon with a low sun angle, you could really start to make out the little back-lit specks in the air up-sun as you descended below about 500 ft agl. It did almost feel like you were descending into a layer of something.
In my current airplane, bug spatters on the windshield mostly appear below 1000 ft, although I get the odd one higher.  The vast majority of the bug spatters on the wings, nose and windshield happen on takeoff and landing.
In the Mooney M20, there is a ram air induction system that supplies unfiltered air to the engine.  The POH only refers to "clear air" on when to use ram air, but if you go on Mooney forums where the question comes up, there seems to be a consensus not to use ram air below 1000 ft due to bugs, and preferably above 5000 ft.
So it's not really a "layer".  If you plotted bug density, you would likely find a parabolic curve with the bulk of the curve below 1000 ft, tapering off above that to effectively zero somewhere above 10000 ft.

The number of bugs on your windshield might also vary due to aerodynamics/ flight regime typical of altitude instead of actual bug density.

Except if you have no windshield.

Yesterday in Oregon I was collecting bugs on my leading edges like a crop duster - at 3500’ AGL.  Low pressure, springtime over agricultural fields and vineyards.
As already answered, most places it is rare to get many bugs over 1000’.  Mylar balloons are almost as common at altitude in California ;-)

Laminar flow isn't necessarily the panacea of design or function; indeed, many aircraft use devices to interrupt laminar flow or energize it.
"Bug layer" is a made-up term, perhaps more applicable to the boundary layer than an imaginary atmospheric one.
I've hit bugs in freefall around 18,000 and have certainly seen birds much higher.  I've hit birds and some large bugs with some frequency above 10,000.'
Icing is perhaps more common and a regular occurrence at higher altitudes.
Some aircraft are much more affected by boundary disturbances, than others.  In the Piaggio P180, even flight through fair weather cumulous clouds at low flight altitudes has a noticeable effect. With the autopilot engaged, it's nearly imperceptible, but when hand flying,  requires movement of the control column back and forth, and the change is quite apparent.  Likewise, bugs or other interruptions of the boundary layer, including rain or any ice, makes a significant difference in control, speed, and climb performance.

ice is melted with de-ice, no?

In flight, generally no.  There are chemical systems in flight such as the general aviation TKS system, and some aircraft that use alcohol for the propeller or radome (eg, Learjet), but most aircraft, no.  Deice and anti ice fluids are used on the ground.  In flight, anti-ice is typically heated leading edges of wings and horizontal stabilizers, and in some aircraft, inflatable boots along the leading edge of flying surfaces, for some arcraft.  Ice can build faster than it can be removed, however, and can disrupt airflow.


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Why is it harder to maintain laminar flow on larger aircraft?
I am told that on larger aircraft, suction will be needed to maintain laminar flow because of the larger leading edge radius. Why so? Is it because the air has to travel over a longer distance on the skin, so slows more, and thus piles up toward the trailing edge, causing adverse pressure?

The flow path is longer, so transition happens relatively earlier. Also, wing sweep and gaps from leading edge devices will quickly trigger transition. The LE radius is not to blame.

"relatively earlier" as in percentage of chord?

wait why would LE radius not have an effect? since the LE slows things down/causes an adverse pressure gradient?

Every nose has a stagnation point, regardless of radius. The larger radius only makes the suction peak near the lading edge at higher angles of attack less peaky.

Sort of interesting that very slow, lower Reynolds number airfoils, such as models, actually have to trip the airflow to form a boundary layer.
Surface friction causes the air near the wing surface to move at the wings speed.  In a wind tunnel this is slower than the free stream, in the air, the moving wing speeds up the air around it.  In relative terms, the effect is the same.
Boundary layers are draggier than laminar flow, but much less draggier than turbulent flow.  The boundary layer is like "grease", helping reduce drag and create lift.
At higher subsonic Mach numbers, early efforts at a "laminar flow wing"  for long range escort fighters  resulted in something resembling an airfoil flown backwards, with a sharp leading edge, a very gently increasing thickness (to keep laminar flow), maximum thickness further back towards the trailing edge, and a much shorter taper to the trailing edge.
This is reminiscent of the blunt "boat tail" approach being tried these days, also found in bullets.
Though successful, low speed handling and stall characteristics led to some needed modifications.
But chord length, and airspeed, as well as airfoil type and surface smoothness, are very important parameters in predicting airflow around a wing.  Airfoil Tools, on the web, is a great place to begin to explore airfoil types for various applications.  Particularly of interest are Coefficient of Lift vs AOA, and Lift/Drag ratio.

when you write "boundary layer" you sure mean "turbulent boundary layer". There is always a boundary layer, and it is either laminar or turbulent.

Thats what i was wondering after reading this answer. so far as i knew, there are laminar boundary layers - the best kind - then turbulent. The only thing that goes beyond that is separated layer, where flow at skin is moving upwind. Am i correct? what's going on in this answer?

@ABJX you are correct.

@ABJX if only we could see the actual flow, "laminar flow" would be near impossible near the surface, where air struck by the wing surface is accelerated and deflected.  Terminology aside, flow separation is what we do not want.  Much depends on Reynolds number, and Mach number as well.  As Peter pointed out in his link, laminar flow is a reality for the underside of the wing, and supercritical design sadly parts with camber that serves so well at higher subsonic Mach numbers.  When I began studying airfoils, I thought battleships had it wrong.  Now, perhaps not!


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Why are inlet vanes used instead of variable geometry inlets to prevent compressor surge/stall?
Jet engine core axial compressors (that of subsonic turbofans anyway) are designed to keep air flowing the right way at max RPM(or close to it). Below that, the airflow slows more and more, and becomes unstable.
In other words, the compressor can no longer compress the large air volume flowing into it. To prevent this, variable inlet vanes are used to reduce the airflow volume into the compressor at low RPM.
My question is: why was this method chosen? When you're running the engine at low power/RPM, the closed vanes will create drag by presenting their blunt face to the air. So why not use a streamlined variable geometry inlet which can forespill the excess air with less drag? such as the following:

TOP: low RPM, a sliding annular door narrows the annular gap
BOTTOM: high RPM, the door slides back, widening the annular gap

What you propose is used as a dust filter for helicopter engines, but even there the plug is not variable. Instead, air spills over the rounded edges of the intake if less is needed.
Intake vanes never close to a point where they would block the intake; instead, they direct the flow such that it hits the compressor blades at the desired angle of attack. They can be moveable to adjust flow to a range of compressor speeds but they never block the flow. Flow control is simply achieved by pumping more or less air into the compressor by means of varying compressor speed. If less air is pumped, more ram pressure builds up in the intake and forces more air to spill over the intake lips. If no separation occurs, this causes little drag and is preferred over something that adds more mass and complexity.
A variable intake body similar to what you propose is only used on supersonic intakes when the shock originating from the tip of such a spike needs to be positioned close to the intake lips in order to achieve uniform flow conditions at the intake face and high precompression efficiency.

If their function is AoA, then why are they ever variable? As I understand, this angle of attack control is needed prevent compressor stall at high airspeed and low RPM. Fixed blades would do the job fine, it would appear. What am I missing?

@ABJX: See here for a list of good answers. In short: With variable stators, the engine has a wider off-design range, AoA adjustment helps to prevent compressor stall at low RPM and reacts faster to throttle changes. Axial flow speed drops faster than circumferential blade speed at lower RPM, so AoA would quickly increase beyond the stall point.

Sorry, by blades I meant inlet vanes

from that link the pressure falls at the discharge of the compressor causing the velocity in discharge (HP ) stages to increase,  which can choke these stages. This causes the front or LP stage of the compressor to increase in loading and can stall these stage causing compressor surge (a violent aerodynamic compressor instability) at such low speed.  Closing variable stator vanes in the front stages will unload these stages thus preventing the compressor from stalling in other words, axial flow is too slow, which is what I said. And yes, I understand that slow axial flow can raise AoA

this too seems to point at there being too much air to handle?


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Can the wing stagnation point be bled off and used for laminar flow control?
The stagnation point at the leading edge contains slow-moving, slightly compressed air. This sounds ideal for laminar flow control. This slow air will create little losses in ducts, and it's pressure allows it to be blown out of slots to reenergize the boundary layer, while at the same time removing contamination - a MAJOR issue for laminar flow today. Is this true?

Don't accept an answer immediatey. Better to wait until more answers have collected and errors have been pointed out and corrected.

No.
Slow-moving also means low mass flow. In order to re-energize the boundary layer, the air has to move fast, meaning high mass flow. Also, speed means reduced pressure, so the stagnation pressure is not sufficient to produce that desired high-energy flow.
Also, contamination (bugs, ice) accretes close to the stagnation point. Ideally, in order to remove contamination, blowing would have to happen across the stagnation point. For that, blowing pressure needs to be higher than stagnation pressure.
The only use of stagnation pressure for boundary layer control is for tripping by blowing perpendicularly to the surface into the boundary layer. This works with low mass flows already.


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Why does the A320 still have this old fashioned nose?
Another answer stated that Airbus was less focused on just brand image, and more on aerodynamics, explaining the more rounded nose of it's 320 vs the 737's. But it still has flat cockpit panes which very conspicously NOT flush with the nose.
The same cannot be said of the A220, which is not much smaller, or even of the A350 which Airbus themselves developed.

You should link the question, not just refer to "another answer".

https://farm7.static.flickr.com/6034/6312829920_819784b2e2.jpg if anyone wants a picture reference

Don't forget A320 is 40 years old while the other two are brand new!

Cockpit windows that sit flush with the fuselage can save a very small amount of drag, but they require large double curvature window panes. Manufacturing such panes is significantly more expensive and there was limited commercial availability for that product when the A320 was being designed.
You can see a few more examples of flat vs curved windows here: https://www.airliners.net/forum/viewtopic.php?t=754093
Cost and availability are a serious concern, as the A320 was already expensive compared to the competition. Paying more and risking delays if double-curved glass production couldn't be ramped up would not have been worth the savings. Airliners aren't just about performance, the very name "Airbus" hints that they're also a public conveyance.
In short, only the newest airframes get to make that small drag saving. Some older airframes had single curvature windows, which are between flat and double-curvature in complexity.
This has been explored in more detail in a broader question.


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Will wingtip drag always be more for a smaller aircraft, keeping constant aspect ratio?
Sometime ago, I came across an article (I couldn't find it) which stated that wingtip drag was only related to absolute wingspan. That is, for the same wingspan and lift force, varying the chord will not change the wingtip drag. I have also seen an answer on Aviation SE where this appears to be mathematically proven.
So my question is; am I correct in stating that , for a given lift force, wingtip drag only depends on wingspan, not chord, and thus that smaller aircraft are doomed to have more wingtip drag than larger ones?

While that article used imprecise terminology, it was correct.Would the answers to this question help to answer yours? Or this answer? Or this?

Yes, these links helped very much. This question is answered. Thank you very much.


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Why does disk loading mean anything when rotors/propellers are not disks?
"The helicopter rotor is modeled as an infinitely thin disc with an infinite number of blades that induce a constant pressure jump over the disk area and along the axis of rotation."
This is spectacularly not true of common propeller disks. Most of the area in propeller/rotor disk is empty space, through which air flows freely from top to bottom.
So why do we take this mostly empty disk and use it's area in aerodynamic calculations, when it would seem that we should rather be using the area of the blades themselves?
It seems to permit "fooling" disk loading equations by, for example, decreasing the blade chord without increasing the span, but still getting the same results from equations. So, what a I missing?

Which aerodynamics text book are you using as your reference point?  Or even "which helicopter aerodynamics textbook" are you using as a reference.

This feels like a question best answered by: read ahead in your textbook until you come across the blade-element theory, then compare its pros and cons with the disk approach.

I'll be honest. I'm just an enthusiast. I usually use Wikipedia.

What are you missing? It works! Why should we make things overly complicated when simple is enough?

 whatever you do, don't study electronics and electricity ... is it hole flow or electron flow?

The disc model is used wherever generalities of behaviour are studied without the need to know individual blade characteristics. It captures the mathematics of those generalities in a simple way that leaves out extraneous detail which would make the maths impossibly laborious.
Typically the aircraft designer will use the disc model in the design of their airframe, leaving the blade details for the propeller or rotor designer to figure out.

then the statement A low disc loading is a direct indicator of high lift thrust efficiency. doesn't appear to make sense, because the lower disk loading=greater efficiency argument appears to draw on the fact that higher bypass ratio=higher efficiency, but as opposed to bypass fans, which have full disks, propellers and rotors generally have very empty disks, giving less blade area compared to the disk area, meaning that most of the disk is not at all affected by the blades.

Hmm... then again, more blades, more interference between blades. Disk loading takes out the blade count and individual characteristics from the equation, thus enabling simpler calculations, as stated above. Disk loading model does not work for everything, but isa nice useful simplification. For example, e does not actually equal mc^2, but that's close enough for most instances... btw, I don't quite understand how disk loading and bypass ratio are related...

Disc loading and solidity (your "empty/full") are very different things and they are not directly related; the designer has to decide how they will relate, taking several other factors into account - not least existing designs as a baseline. Also propellers, rotors and ducted fans are very different conditions with very different design solutions. There is no space in a comment to elaborate usefully.

"I don't quite understand how disk loading and bypass ratio are related" - there seems a lot that you do not understand. Maybe some reading up on the subject would help, StackExchange does not do taught courses.

The disc approach is just a calculational tool of convenience that contains simplifying assumptions which make the math easier. It works well enough in certain circumstances to justify its use.

What are these "certain circumstances"? It seems to me that the disk idea only makes sense for a fan.

@ABJX What's the difference between a fan and a propeller or rotor?

well, "fan" in the "turbofan" context has propeller blades taking up most or all of the disk area. propellers and rotors have a lot of free disk area.


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Is high propeller pitch inherently inefficient?
For an airfoil, in general, the aerodynamic force is more or less perpendicular to the chord line.
A low-pitched propeller blade has it's chord line more perpendicular to the direction of flight, therefore, more of the force acting on it can have a greater component parallel to the direction of notion, and less perpendicular. If this force is acting forward, then this is efficient. But this may not be possible at high speeds and low rotation velocities. Thus:
A high-pitched blade has it's chord line more parallel to the direction of flight than the above. Therefore, the aerodynamic force acting on it is likely to have a greater component perpendicular and lesser component parallel to the direction of motion. This is always inefficient.
The above manifests itself in the higher-pitched propeller producing more energetic swirl than the lower pitched propeller.
So my conclusion would be that all else being equal, high-pitched propellers are inherently inefficient. Is this correct? Is this what is observed in practice?

Are you talking about fixed pitch props?

@skipper44 any.

Let's imagine a plane fitted with a variable-pitch prop. For a given prop revs and a given forward speed of the airplane, there's an optimum pitch so that the prop blades work with an optimal forward component of the aerodynamic force. The prop revs remain more or less constant, and as the plane forward increases, the pilot has to adjust the pitch to higher and higher values. As the (linear) pitch tends to infinite, the forward component of the blade's aerodynamic force tends to zero. So, yes... The propulsive efficiency of a variable-pitch prop starts to decay beyond a certain pitch value (geometric pitch of about 45º) where it reaches the maximum, and tends to zero as the geometric pitch approaches 90º...

Graph from http://www.epi-eng.com/propeller_technology/selecting_a_propeller.htm

It’s true that there will always be a component of the blade’s thrust that directly opposes the rotation of the propellor, and hopefully a component that propels the aircraft.  However, it would be a mistake to think that the former is wasted, on the contrary it is essential because that force is the source of the energy from which forward thrust is derived - if no energy were required to keep the propeller spinning then there could be no thrust.  Another thing to consider is that steeper pitch propellers have a smaller ratio of blade speed to aircraft speed, and so will tend to be ineffective rather than inherently inefficient.

Mulling this over (and I know here I'm not answering the 'what happens in practice' question), the aerodynamic force generated by each blade can be resolved into two components; one in the plane of the propellor, which applies a load to the engine and one that provides forward thrust to the aircraft.  Mathematically, the magnitudes are given by the sine and cosine of the angle of the aerodynamic force (which will be more or less perpendicular to the chord line).  The energy in each vector is given by the force multiplied by the airspeed in that direction.  Thus a low pitch will result in a...

small retarding load on the ending but a high rotational speed, which is converted to a large thrust force but at a low speed.  Inefficiency would not be caused by a large (or small) ratio of forces; we see that an aerofoil can do this quite efficiently, but more likely by parasitic drag (tip vortices).  It is well established that large propellers are more effective and efficient than small ones.  Surface drag on the propeller will be greatest when the blade speed is large, which suggests that a slow-rotating propellor could be more efficient, and that would require a steeper pitch.

Your conclusion is too simplistic. On aircraft, the difference between a cruise propeller and a climb propeller is rather small. However, in other applications, a low pitch propeller gives great acceleration but the engine will reach max RPM faster, which reduces top-end performance. A high pitch propeller will result in slower acceleration but higher top speeds.
This answer is also admittedly very simplistic -- there is a universe of math and engineering that goes into propeller designs. Sensenich, for example, produces more than 100 fixed-pitch propellers for aircraft, airboats and UAV's. McCauley Propellers produces thousands of constant-speed and fixed-pitch configurations for commercial, military, UAV, and general aviation applications.
If you want to get into the nitty-gritty of this, I have been told this book is one of the propeller design bibles for aircraft applications, "Aircraft Propeller Design" by Fred Weick, published in 1930.


text:
How far back can the FAA go in requiring documentation of IFR currency?
I did an IPC recently and got to thinking how long I needed to retain the proof of re-up. Let's say that a pilot stays IFR current for 10 years after the hiatus requiring an IPC. For whatever reason, the pilot can document staying current during that period, but has lost the IPC endorsement and so cannot document the the chain of currency was founded legitimately.
It certainly would beg belief that the FAA would pursue an administrative action against a pilot who simply cannot document an endorsement a decade prior, but would is not the same as could.
How far back can the FAA demand a chain of proof?


text:
What is the logic behind requiring an instructor endorsement for only certain FAA Knowledge Tests?
Reference: https://www.faa.gov/training_testing/testing/testing_matrix
Some tests, such as for the PPL, require an instructor's endorsement. Others, such as for the CFI, do not. Is there any logic behind this, or is this just an artifact of the merge of two historically separate test approaches?

Neither the CFI or ATP written tests require an instructor endorsement.  I would argue that only pilots who are making aviaiton their profession to take those tests and so the failure rate would be lower.  The endorsement is to ensure the applicant is ready to take the test.  The CFI and ATP applicant can self-evaluate if they are ready to take the test.

@wbeard52, I'd upvote that as an answer...

The CFI written exam includes 100 questions of a pool about a 1,000 and most of the questions are common to the Private and Commercial Pilot written exams. If you already passed those there isn't a lot of reason to require an instructor to sign you off to take the CFI test.

Not much logic in requiring you to take it at all, if that's the case!

@757toga So why not post this as your answer? This doesn't change what I said.

@757toga Then the comment belongs below the question, not below an answer.

Agreed that this doesn't particularly answer the question. There's no reason I can see for a CFI ground endorsement at all, ever. But that's not the way the system works, and I won't change it by complaining here. Still, it's interesting that there is this divergence, as anytime there's a divergence there's a reason. And as much as people like to mock regulators, their reasons are not usually crappy and ill-thought-out. When the rationale is crappy it's almost always as a result of being stuck between an unstoppable force and an immovable object. What, then, are those forces and objects?


text:
Does a BasicMed certificate need to be printed to be valid?
I understand the logic behind the FAA wanting a physical card on the pilot's person-- it's so an FAA rep can revoke the license on the spot. However, this logic doesn't apply in the slightest to the medical.
Do the Certificate and the Checklist, both required for BasicMed validity, have to be printed out or is it legal to fly with them in digital form, e.g. on a phone?


text:
What was the FAA's official rationale for the ADSB mandate?
This might have been obvious at the time, but I can't find any articles that discuss the FAA's official standpoint on requiring all GA planes to equip. There was and still is a hefty cost to ADSB compliance, which AFAIK means the FAA would have to justify in a NPRM why the mandate produced more value than it cost.
I'm curious what that rationale was.

The FAA puts the rationale for rule changes in the Federal Register as part of the Notice of Proposed Rulemaking (NPRM).  The NPRM for ADS-B Out was published on 10/05/2007. Link: Here
It's 26 pages long. I'll include some of the key points here:
From the FAA's NEXGEN vision:

The third element—performance-based surveillance—relies on technology
that permits knowing the exact location of other aircraft in the air
and of other aircraft and ground vehicles on the airport surface. The
aviation community's experience with ADS-B, which periodically
broadcasts an aircraft's location—both horizontal and vertical
position and horizontal and vertical velocity—will lead directly to
the performance requirements. When displayed in the cockpit,
information obtained through ADS-B greatly improves situational
awareness in the en route segment, in the terminal area during
approaches, and on the airport surface. For additional information on
ADS-B activities, see Section VI, FAA Experience with ADS-B later in
the preamble.
This rulemaking is important because ADS-B is an essential NextGen
building block. Improving surveillance requires advanced onboard
equipment with backup capability. Most, if not all, of the
surveillance capability as well as the navigation and communications
capabilities should be onboard the aircraft so the required
capabilities will go wherever the aircraft goes. As part of the
rulemaking effort, the FAA established an Aviation Rulemaking
Committee under Order 1110.147. This committee has been chartered to
deliver a report on how to optimize operational benefits of the ADS-B
system and to provide recommendations to the FAA on the rulemaking
after the NPRM is published. The scope of the ARC membership is
designed to provide the widest range of inputs into the development of
the NextGen strategy. The FAA will put the ARC recommendations in the
docket established for this rulemaking.
It is this combination of onboard capability and performance
expectations that will enable aircraft in the future to fly safely and
efficiently despite ever-increasing demands on the airspace.

About the limitations of radar:

To accommodate the projected level of traffic without increasing
delay, more comprehensive surveillance in the NAS, including more
radar sites in certain areas, would be necessary. Even if more radar
sites were commissioned, however, there are many areas in which radar
coverage is not feasible, either geographically (e.g., mountainous
areas) or in a cost-effective manner (e.g., remote areas).
Furthermore, simply increasing the number of radars in the NAS does
not solve the inherent limitation of radar technology, and would not
allow the FAA to reduce current separation standards.[5]
Consequently, the future of air traffic surveillance cannot be based
solely on the use of radar. Radar technology also lacks the capability
to provide services on the flight deck. However, the FAA is planning
to maintain its current network of primary radars, and expects to be
able to reduce a percentage of its secondary radars. This NPRM does
not propose to reduce primary radar sites.[6]
Instead, this NPRM would transfer future aircraft surveillance to
newer and more advanced onboard avionics that provide more accurate
and timely aircraft information. ADS-B has been identified as the
technology to facilitate that goal.

ADS-B System:

Implementation of an ADS-B system would not completely replace the
primary radar or SSR at this time. In addition, ADS-B does not replace
the requirement for transponders. Transponders are still necessary for
SSR, which is the FAA's backup strategy in case of ADS-B failure. For
more information on the backup strategy, see section IV.C.4, Backup
Surveillance Strategy.
The performance requirements for ADS-B avionics proposed in this NPRM
would ensure that the aircraft is broadcasting the requisite
information with the degree of accuracy and integrity necessary for
ATC to use that information for surveillance.[7]
This enhanced surveillance would provide ATC with the enhanced ability
to surveil and separate aircraft so that efficiency and capacity could
increase beyond current levels to meet the predicted demand for ATC
services while continually maintaining safety. Incremental
developments in capacity, efficiency, and air traffic control
procedures based on radar technology cannot accommodate the
anticipated increase in demand for surveillance and separation
services, which could result in delays that would far exceed those
experienced today. Without ADS-B, the increase in demand could result
in increased congestion and the denial of ATC service to some users of
the NAS.
ADS-B technology already has been demonstrated successfully in Alaska
via the Capstone program.[8]
In Alaska, radar coverage is either very limited or non-existent.
ADS-B provides a level of surveillance performance that previously did
not exist and has resulted in increases in both efficiency and
capacity.
“ADS-B Out” refers to an appropriately equipped aircraft's
broadcasting of various aircraft information. “ADS-B In” refers to an
appropriately equipped aircraft's ability to receive another
aircraft's ADS-B Out information. This proposal only seeks to require
ADS-B Out; the FAA is not proposing to require ADS-B In at this
time.[9]

Very nice. As a way to enhance my searching skills, could you let me know how'd managed to find that so quickly?

It would be good to summarize the answer here quoting from the official source, otherwise this answer is your answer is in another castle

@Jamiec Point made. Done.

Regarding the search. The first point is understanding the rulemaking process. That came from working in avionics for 25 years.  That process normally starts with a Notice of Proposed Rule Making (NPRM) being published in the Federal Register, which is an official government publication. In it, they propose the rule, why it's needed, and solicit input. At a later point they publish a Final Rule including a discussion of the comments received on the NPRM and changes made. So I just Googled "ADS-B Notice of Proposed Rule Making". I then looked thru the Federal Register links for the first one.


text:
What is the range of ratios between optimal propeller efficiency and true propeller efficiency?
In https://aviation.stackexchange.com/a/15383/20394, the optimal prop efficiency is given as:
$$\eta^{opt}_{Prop} = \frac{1 - \frac{\Delta v\cdot \left({v_{\infty} + \frac{\Delta v}{2}}\right)}{d_P^2\cdot\Omega^2}}{1 + \frac{\Delta v}{2\cdot v_{\infty}}}$$
If we think about $\eta_{Prop}$ as $k\eta^{opt}_{Prop}$, where $0\leq k \leq 1$, what would be the typical range of $k$?
Furthermore:

Can we draw general conclusions about $k$ for various props, based on material and the modernity of its construction/materials/airfoil?
How good of a fit is the $ \eta_{Prop} = k\eta^{opt}_{Prop}$ model for the true prop efficiency?

P.S. I am making a distinction between propeller efficiency and propulsive efficiency. This question is only about the former. C.f. Propeller Efficiency calculation different by two methods

My XROTOR.DOC file from October 1988 reads:
<--- snip --->
Like CL, profile CD is defined piecewise in alpha such that in the unstalled
region, CD has a quadratic dependence on CL (after Prandtl-Glauert scaling)
and a power-law dependence on Reynolds number as follows.

            |                   2|           f
     CD  =  |CD  +  b (CL  - CL) | (Re/Re   )
            |  o         o       |       ref
where

CD    =  minimum drag coefficient        ; Fortran name:  ( CDMIN )
  o
CL    =  CL at which CD = CD                              ( CLDMIN )
  o                         o
b     =  CL weighting coefficient  d(CD)/d(CL**2)         ( DCDCL2 )

Re    =  Reynolds Number at which CD formula applies      ( REREF )
  ref
f     =  Reynolds Number scaling exponent.                ( REXP )
         Typically:
          f = -0.1 to -0.2  for high-Re turbulent flow
          f = -0.3 to -0.5  for low-Re airfoils

Only the above parameters are supplied by the user, and the program
figures out the CD(alpha) function in SUBROUTINE CDCALC.

<--- snip --->
So we run XROTOR once with realistic drag data and once again with the same propeller but all drag coefficients set to zero. Since I last used XROTOR in the last millennium, I have only my old MacOS version which I run in Sheepshaver. It took me a while to get familiar again, and I took Mark Drela's Daedalus propeller file as the input. Copy-Paste from the old to the new OS doesn't work, so I have pasted screenshots here:

Next, I use the AERO submenu to set the drag parameters and redo the calculation at the same RPM:

What a surprise, without airfoil drag our efficiency equals the ideal efficiency. Looks like I did everything right to isolate viscous effects. The difference is a jump in efficiency from 0.886 to 0.94, a bit more than 5%. This is for a very lowly loaded prop at a very low Reynolds number, so the viscous effect is rather large. Expect less for larger props at higher speed, so your factor k will range between 0.943 and 0.984. The last number is valid for a 6-bladed prop driven by an AE2100 engine, the input file of which I also had at hand.

Erhagherd, the level of dedication to make this answer! You deserve all the upvotes I can muster for this one.


text:
What justifies a biplane over a braced monoplane?
Biplanes became the first airplanes to enter service due to their structurally efficient design being more suitable for the weak materials then. This efficiency comes from the two wings acting as the chords of a very deep truss.
Another structurally favorable factor was the reduced wingspan and thus moment allowed by distributing lift across stacked wings.
Yet another weight saver was the reduction of wing loading, which enabled thus use of less powerful engines for slower flight.
However, the extra wing only produces 20% more lift with 100% parasitic drag increase. Ironically, this interference drag is even more severe at the lower speeds which biplanes are optimised for.
A braced monoplane has most of the structural advantages of a biplane, with less of the aerodynamic disadvantages. So it would seem to be a more logical design than a biplane, given the same materials and engine.
Yet biplanes were built in large numbers. Why?

BTW I honestly question the lower wing gives only 20% more lift.  It may be more, if set up properly, the lower wing puts the upper wing in perpetual ground effect.  It would be a great study, but, as Peter says, the upper wing does most of the lifting.

Guy Inchbald's answer was pretty accurate.

@AbdullahisnotanAmalekite just to understand why the bounty: what is missing in the answers that you got so far?

Back in the pioneer days both types competed fiercely. The braced monoplane has similar structural depth to the braced biplane but less than twice the total wing area, so many argued in its favour. On the other hand, the shallower angle of its bracing wires introduced greater compressive forces in the wing spars, so they had to be strengthened and that added weight.
The monoplane suffered its first drawback around 1911 when, first in France and soon after in Britain, examples started falling out of the sky and for a time they were banned from military service. The French problem turned out to be a mistaken design assumption leading to over-thin bracing wires, the British to be the Ministry-specified cowling fasteners coming adrift. They were soon flying again, but the damage to their reputation was done.
Nevertheless, a few types were still in service when war broke out, especially with Germany. The Etrich Taube and Fokker Eindecker helped to restore their reputation, especially when Fokker developed interrupter gear for the machine guns, unleashing the Fokker scourge. But this would prove the high point of the braced monoplane's career.
The advantages of the biplane you mention - high strength and short span - proved to give greater manoeuvrability, stiffness and damage tolerance when it mattered, and the monoplane quickly became obsolete.
The biplane remained dominant until engine powers rose sufficiently to take planes beyond 200 mph or so, where the greater weight of the unbraced or cantilever monoplane was a price worth paying for cleaner aerodynamics.

The biplane lost its dominance already during WW I, at least in Germany. Most new types were monoplanes: Fokker DVIII, all new Junkers, Dornier or Rohrbach (Zeppelin Staaken) designs, you name it. New biplanes were mostly designed for aerobatics and pilot training.

@PeterKämpf That is a false picture full of mistakes. The parasol monoplane mounted only a modest challenge, and that mainly in France. The most significant Dornier was the D.1, a cantilever biplane, Junkers had built cantilever monoplanes throughout but only in penny numbers, Rohrbach were a postwar company; biplane houses such as Fokker and Albatros dominated German skies to the end. French fighters would serve your case better, but companies such as Lioré et Olivier long continued the biplane tradition in other sectors.

Rohrbach worked at Zeppelin in Staaken. That's why it included the subtle hint that you close to overlook. The Dornier D.1 was partially responsible for the demise of biplanes. None of the many Albatros biplane types after the D.V were ordered in numbers. Fokker changed to monoplanes in 1918. Dornier and Rohrbach designed the planes you call postwar in 1918. Should I continue or do you see where the misconceptions are?

Zeppelin Staaken did not build Rohrbach's monoplane until the war was over, and one obscure prototype doth not a revolution make. The Fokker D.VIII monoplane was smaller and lighter than the D.VII biplane but no faster, and the company went back to biplanes for what little was left of the war. Yes, I think we can all see where the misconceptions lie. The simple fact of history is that monoplanes remained a curiosity until engine powers rose to justify the structural inefficiency of the cantilever.

Four elements of the aviation matrix compete, as always: structures, aerodynamics, engine power, and stability/control. The very first powered flight of the Wright brothers took place in a biplane with a 12 hp engine, and was hard to control. It took almost a year after the (brilliant) first flight before they could execute a full 360 degree turn, and six of their aircraft delivered to the US Army crashed. From the wiki:

Army accidents
In 1912–1913 a series of fatal crashes of Wright airplanes bought by the U.S. Army called into question their safety and design. The death toll reached 11 by 1913, half of them in the Wright model C. All six model C Army airplanes crashed.

How to make aeroplanes stable and easy to control? One solution was implemented by the young Anthony Fokker: simply eliminate controllability in the roll axis. `In the pic below from the Wiki article about the Fokker Spin, Anthony sits behind the engine structure, and the two wing halves can be distinguished supported by a spider web of bracing wires (Spin translates as Spider). Large wings => low wing loading!

The two wing halves can be seen to tip upwards, providing passive roll stability. No ailerons, the flight direction was commanded by the rudder only. Flying slowly in a straight line, turning gently with rudder only. And then WW1 started. At first aeroplanes were just observers, then came to be pursued by enemy aeroplanes to eliminate the flying artillery scouts. Airborne manoeuvring was now applying large loads on the structure, and the monoplane consisting of two wing halves held in place by wires could not compete anymore. For one, more wires were required, creating way more drag.
So the solution was the inherent strength and stiffness of the biplane structure. First the rectangle was crossed out by wires, later by more aerodynamic wooden bracing. At the end of WW1 the most feared plane was the Fokker DVII, which was particularly mentioned in the peace treaty.
Does the lower wing of a biplane account for only 20% of the lift? For sure if its area is 20% of total wing area. Aerodynamically an unbraced monoplane is superior, structurally a biplane is simpler. If the function of the plane is transport for low costs, a single wing with a large aspect ratio is the design goal. A small aerobatic plane might still be a biplane though (Pitts Special, from the wiki).

I fail to see how a Mono wing requires more wires.

A modern mono win does not, because of the uninterrupted construction through the hull. In 1911 it did: the two wing halves can flap like a bird if not wired.

So they made the mistake of having the wing split by the cockpit?

Well that was the technology at the time. The structurally uninterrupted monoplane came later - I had actually posted a question about this, the Junkers F13 of 1919 seems to be the first unbraced monoplane


text:
Why aren't isentropic noses and tails used in supersonic aircraft?
An inside out paraboloid would theoretically produce isentropic compression for an inlet cone. But this question is about nose and tail cones on fuselages, and leading and trailing edges of wings.
An isentropic nose would preserve pressure in the region surrounding the aircraft, so that more of it may be recovered at the tail. This effect, if course, may not be very significant. It would also reduce aerodynamic heating.
An isentropic tail would enable better pressure recovery following the expansion fab, and thus reduce drag.
So, why are these shapes not used?

whatever physical process increases enthropy, how can you have "isoenthropic" (i.e. no-enthropy changing) anything?

@Federico give me a break hehe


text:
Does altitude affect glide ratio?
Lower altitude means higher air density, which means more liff, which means a glider can maintain a given glide slope at lower airspeed.
As a rule, aircraft tend to have higher peak L/Ds at low airspeeds, due to a reduced parasitic drag coefficient.
So, for a given design, I would expect the best glide airspeed to increase with altitude, and best L/D to decrease, and best glide slope to steepen.

The ratios of L/D and glide slope ought to remain constant, but TAS will be higher.  (Presuming best glide IAS is constant)

That's an answer, although you may need to add some flowery language to bulk it out a bit.

@John K, yeah, I thought about puffing it up and making it an answer, but as soon as I bother taking the time some aero major will chime in with a bunch of formulas and upstage me!

Keep in mind that there is very little difference in lift force (as opposed to lift coefficient) between a plane's optimum glide ratio, and a greatly degraded glide ratio.  Nearly all the variation in Lift/Drag ratio is due to variation in Drag force, not Lift force.  Lift  = Weight * cosine (glide angle).

Also, suggestion-- clarify whether you are talking about IAS or TAS or -?

PS don't forget about Reynold's number--

Yes, altitude does affect glide ratio, but only slightly.
With increasing altitude, air temperature drops and with it the Reynolds number of the flow. This is mitigated, but not completely offset, by the increase in true air speed, such that the Reynolds number of an aircraft in the tropopause is less than half of the Reynolds number at sea level.
Since the Reynolds number is the ratio of inertial to viscous forces, a lower value signifies more viscous effects and, hence, drag for the same lift.

The L/D ratio will be the same at any altitude as the Lift and drag are related to the dynamic pressure experienced by the aircraft, which is a function of Indicated Airspeed, however, the True airspeed will be higher at higher altitudes. This means that a glider descending from 10000 feet to 9000 feet will cover the same distance over the ground as if it descended from 2000 ft to 1000ft, but it would take less time as its ground speed would be higher. This would give the impression that the glider had a higher performance at high altitudes.

Altitude does not affect the glide ratio, since air density cancels out in the  ratio L/D.
L/D is numerically the same as glide ratio (in conditions of zero wind). In this figure of a gliding plane, taken from 'The simple science of flight', by Henk Tennekes, it's  clear that the triangle of forces L-D and the triangle of velocities U-w are similar. Hence, the ratios L/D and U/w are the same...

Seems good, but- what about Reynold's number?


text:
Making sailplane go faster
Sailplanes have much higher L/D than jetliners. But they are three times slower.
Is it practical to make a vanilla sailplane glide at 600mph by lifting it to a ridiculous altitude? If so, what would be the quantitive effect on L/D?

Check out Perlan 1 and Perlan 2 high altitude glider projects. According to Wikipedia, Perlan 2 has a $V_{ne}$ of 434 mph (TAS). When at it's service ceiling of 90k feet, it is designed to fly at around around 400 mph.

aerodynamic flutter has entered the chat

Use this calculator to work it out. If your glider reaches it's L/D max at 70 kt indicated, pretty typical, and you tow it to 50000 ft, its max L/D will still be 70 kt indicated, but the TAS will be 140 kt.
To get to 600kt TAS, you'd be flying it at 300 kt indicated at 50000 ft. Well past the typical glider's Vne of, say, 120kt. So you would be diving vertically, and it would likely come apart first.
At 90000 ft, at 70kts indicated, TAS is 196kt. To get to 600 kt, you'd need to be at over 200kt indicated, still well exceeding the Vne of a typical glider. Best you could hope for is 350-400 kt and not exceed Vne, and that would still be a vertical dive. If you're gliding for best range with a Max L/D at 70kt indicated, 196 kt TAS is it.


text:
How to determine Center of Gravity of the airplane parts?
I am building a small-medium sized airplane.
Is it possible to calculate "C.G" of each part of the airplane? If yes, How?
Parts: Wings, Tail, Fuselage, Propeller(?)

While it would be possible to calculate the centre of gravity of each component of the aircraft, I'm not sure what the point of that would be. Do you want to calculate the position of the centre of gravity of the complete aircraft?

I do want to calculate the position of CG of the complete aircraft..
One thing i want to point out is... I am building the plane from scratch for the college project (not using standard parts, that's why i need to mathematically calculate all of this) i'm having problems finding formulas and resources to calculate the C.G of each parts of the plane.
I'm also using this information to build wing, tail and propeller in relationship to the fuselage to make the thing fly with stability.

Calculating the CG of either the complete aircraft or the major components should be relatively simple if you are using CAD/CAM

If you are building a light single engine low wing monoplane, the exact position of the CG will vary depending on who is flying it, the fuel load and payload. The shorthand method of getting the CG in about the right place is to arrange the major weight components in the fuselage so as to place the CG of the fuselage over the center of lift of the wing.

Let’s start with a definition from Airplane Flying Handbook:

Center of gravity (CG).
The point at which an airplane would balance if it were possible to
  suspend it at that point. It is the mass center of the airplane, or
  the theoretical point at which the entire weight of the airplane is
  assumed to be concentrated. It may be expressed in inches from the
  reference datum, or in percent of mean aerodynamic chord (MAC). The
  location depends on the distribution of weight in the airplane.

Your center of gravity calculations will start out by deciding where you want to place the datum. Typically it is placed at the firewall, but sometimes it is at the tip of the spinner.
For every piece on the airplane you measure the distance from the datum. That is called the arm. For my Cherokee the front seats are at 37 inches from the firewall. Let’s say that the seats weigh 20 lbs. The moment is 20*37=740. There are two of them, so total moment is 1480. Do this for every item on the airplane and then add them up.
Divide the total moment by the total weight to get the center of gravity.
Lots of things are fairly easy to estimate: engine, prop, battery, seats, cowl, tail have easy to calculate centers of mass.
Where it gets complicated in your case is that some of the parts of the airplane can’t be treated as residing at a point. You would need to balance them to find their individual center of gravity. Or you could break them down into parts and add up the parts. For example, you could calculate the weight the last two feet of the fuselage and use the middle of that as its ARM. Then the next two feet, etc. Same with the wings.
If you are using standard parts there may already be CG information from other builders or the supplier.
This will give you an idea as to where the CG is but when you are done you need to weight the airplane and calculate the true CG. This article goes into detail on the method and calculations.


text:
Ground Power KVA when on stand
How much power from a GPU would a jet, say a Boeing 737, require when boarding passengers or just sat idle on the air bridge?
I’m guessing the Aircraft would only consume the full 90kVA ground power rating when starting the aircraft’s APU?
Thanks,
Chris

From the amount of noise they make, I'd say tons of power!  Nothing more irritating than the high pitched whine of an APU or GPU.  Except maybe neighborhood leaf blowers running for hours on end...

Depends, are they using air conditioning, and how much? This is the main power consumption on the ground, other than starting; some airports have ground conditioned air. I'd be surprised by more than 10-20kVA in average conditions


text:
What does mean 'GS' in RNAV (GPS) Chart?

Hello.
I have a question about 'GS' in my RNAV(GPS) Approach Chart.
Usually, (Including PAPI, VASI whatever...) Normal Glide Descent Angle means 3.00 degree.
but some 'GS' says 3.07, 3.04, 3.09 even 7.75 in KSBS airport. (Picture above)
In this case, What does mean 'GS' exact ?
As far as I know, it's 'GS' means draw an imaginary line extend from Touch-Down Foot marker (1,000ft marker) to FAF for Non-Precision Approach.
(Even do I Step-Down Approach ? I also wonder.)
but someone says, it's a line extend from 1,000ft marker to VDP,
and another one says, it's a Normal Descent Rate when PAPI indicates 2white-2red. (Exact Center-line of PAPI Glide Slope. I know approximately 0.4 of marginal has exist on PAPI for 2White-2Red (2.81-3.19)
If the last one's word is true, an airport have a 7.75 GS, like a KSBS Airport, then the PAPI of KSBS airport indicates 2white-2red when I follow the 7.75 GS ?
Waiting for your answers. Thank you.
BEST

Are you asking about the specific term 'GS'?  I don't see that on the chart in your question.
The angle symbol and the 7.75 are explained as a Visual Descent Angle in the Chart Users Guide.

A VDA and TCH may be published on non-precision approaches. For Copter approach procedures, a Heliport Crossing
  Height (HCH) will be depicted in place of the TCH. The VDA is strictly advisory and provides a means to establish a stabilized descent to the MDA.

So it is an aid for maintaining a stabilized approach that should put you at the MDA at the correct point.

If the last one's word is true, an airport have a 7.75 GS, like a KSBS Airport, then the PAPI of KSBS airport indicates 2white-2red when I follow the 7.75 GS ?

No.  In fact it's called out on the chart: "VGSI (Visual Glide Slope Indicator) and descent angles not coincident".  So they do not match.  You can look in the chart supplement for the airport and see what the actual angle is.  For KSBS runway 32, the Glide Angle (GA) is 4 degrees.

RWY 14–32: H4452X100 (ASPH–GRVD) S–50, D–60 HIRL
RWY 14: Rgt tfc.
RWY 32: REIL. PAPI(P2L)—GA 4.0º TCH 40´. Thld dsplcd 600´.

GOT IT !! and I might be little confused about 'GS' and 'VDA' in context....... and.... yes, yes, you're right. The words which I used 'GS' in context, was 'VDA' actually. Thanks :D. And I have a one more question. can you let me know 'VGSI and descent angles not coincident' ? As far as I know, All PAPI of USA indicates 2W2R on 3.0 degree path. but as you know, For KSBS R32 says 'GA' is 4 degree. Then In this case, If I descent with 4.0 Degree, the VGSI Indicates 2W2R in lieu of 4W? or just 4W?..... Then How can I find I'm on gilde path(4.0) or not ?

3 Degrees is nominal for PAPI, not required.  For obstacle clearance, it may go up to 4 degrees.  See https://www.faa.gov/documentLibrary/media/Order/FINAL%20FAA%20Order%206850.2B.pdf Page 5-7, note 1


text:
Does lift-to-drag ratio asymptote to 4 (or 6) at high Mach numbers?
The Kuchemann equations for L/D at high mach numbers, approximately verified by wind tunnel tests, were:
$$\left(\frac{L}{D}\right)_{max} = \frac{4\cdot(M+3)}{M}$$
and, for wave riders
$$\left(\frac{L}{D}\right)_{max} = \frac{6\cdot(M+2)}{M}$$
As you can see, these equations asymptote to 4 and 6 respectively.
If my understanding of the matter is correct, at ultra high Mach numbers, air begins to behave more like rays of radiation. Hence, L/D would simply be:
L/D = 1/tanA

where A is angle of attack.
Hence, the corresponding angles of attack for the two asymptotes would be 14 degrees and 7.5 degrees, respectively.
Is there any physical evidence for the above asymptotes (and perhaps the AoA values)?

My source is STUDIES OF HIGH LIFT/DRAG RATIO HYPERSONIC CONFIGURATIONS by John V. Becker, to be found online here. It is from 1964, so more than 50 years old, but since a lot of research had been performed already before that date, it might still be relevant.
In short: It depends. On thickness ratio and Reynolds number, for example, as can be seen in this plot for flat plates from the above linked report:

Note the grey areas noting the maximum for glide rsp. reentry vehicles. Their performance is roughly half as good as that of a flat plate.
Most of the report uses the volumetric efficiency parameter $\frac{V^⅔}{S}$ for comparisons. This is the volume $V$ to the power of ⅔ divided by the reference area $S$. Below is a carpet plot for a variety of wedge shapes, valid for Mach 6.8 (same source as above).

An infinitely thin wing will achieve an L/D of 6.5, irrespective of its leading edge sweep, and a more realistic t/c of 0.09 will not get above an L/D of 5.5 at 85° sweep, dropping below 5 for 70° sweep.
It seems the Küchemann approximation holds up well for single-digit Mach numbers but becomes too optimistic above this speed. And impact theory alone is not sufficient to predict L/D with any confidence.
The plot below explains why the Shuttle is a low wing configuration: Above Mach 10 it is cleary advantageous to put the flat side on the bottom:

The index FB represents the flat bottom version of the vehicle.

Curiously, when we plug some numbers, L/D comparison of the "waverider" get better and better as Mach increases:

M
$\frac{4 \cdot (M+3)}{M}$
$\frac{6 \cdot (M+2)}{M}$

1
16
18

2
10
12

3
8
10

4
7
9

5
6.4
8.4

This confirms that supersonic aircraft do better as "bottom lifters", essentially an extrapolation of supercritical wing design.
The XB-70 prototypes took advantage of this by folding its wingtips down in flight, and riding its own shock wave.
Supersonic wings are very thin and diamond shaped.  Add a little AOA
and one can see a parallelogram, with very little "classical" top lift being produced.
In the supersonic realm, the same lift with less drag is achieved with the waverider.

It should be no surprise that hypersonic lift is mostly created on the lower surface. With the low density in the hypersonic flight altitude, the pressure difference on the top surface cannot exceed that between a vacuum and atmospheric pressure, while on the bottom the pressure can add up to multiples of  atmospheric pressure.

Thanks Peter Kampf and ROIMaison


text:
Why do sailplanes have such large tails?
Since a sailplane is focused on performance - what with tricky laminar flow airfoils and all.
Eo why this

Instead of this

It looks to me like that long tail is needed to balance the cockpit. Buy then why not just put the wings beside the cockpit?
Even if we cannot use a flying wing, why not s least a shortened tail? Why this weight and drag penalty?

The Genesis isn't really a "shortened tail" regular glider.  It's a flying wing glider, designed by Jim Marske, who wanted to keep it a pure flying wing, but was browbeaten by his financial backers into putting a little horizontal trimming surface at the top of the fin. They eventually threatened to pull their backing if it didn't have a surface there, as they thought it would limit sales without it.  Its horizontal "tail" is a redundant appendage stuck there to keep his partners happy.

@qqjkztd that thing is a flying outrage! Also, I would think it needs a wash-in

@qqjkztd I think a better comparison would be glider A with a long boom/small surfaces and glider B with a short boom large surfaces. The Genesis is really kind of an oddball with a unique story.

@Abdullah Marke's designs, incl the Pioneers and the Monarch, have very nice flying qualities as flying wings go. The forward sweep is modest and is offset by generous fin area and give decent yaw stability by glider standards.  They are effectively spin proof because the forward sweep inhibits tip stall and the inboard wing, where the elevator is, cannot be brought to stalling AOA because the elevator unloads as soon as the flow at the TE starts to separate. As a result, you don't need wash-in, which promotes tip stalls and spins.

I had trouble in KSP until I looked at just how stupidly large the tail on a B-29 is (more than doubles the height of the aircraft). If you can go fast enough, the only thing you need is a tail. However big you think it should be, double it.

Keep in mind that not all sailplanes are designed for extremely light weight.  Decreasing the weight also decreases the minimum sink rate, but doesn't help the still-air glide ratio, and actually hurts the glide ratio against a headwind.
Even in sailplanes that are designed for very light weight, designers have sometimes thought it worthwhile to employ a rather large vertical fin.  A good example of this is the Carbon Dragon.
The basic purpose of a vertical fin is to prevent the aircraft from flying sideways through the air, especially when the pilot makes an aileron input, which tends to create adverse yaw.  Adverse yaw tends to be most pronounced in long-spanned, slow-flying aircraft.
Flying sideways is never efficient.
However, not all sailplane designers have always followed this philosophy.  The Scheibe Bergfalke is an example of a glider with a rather small vertical fin.  (Look closely at photos of this glider and you'll see that more than 75% of what appears to be the "tail" is actually the rudder, not a fixed vertical fin.)  When flown with the pilot's feet off the rudder pedals, it tends to be rather uncooperative in response to aileron inputs, being quite happy to fly sideways through the air rather than actually turning.
Keep in mind also that a given desired "tail volume"-- the product of tail area times moment-arm-- can be achieved with less drag with smaller tail on a long slender boom (rear fuselage), than with a larger tail on a shorter fuselage.  You've really asked two separate questions here-- "Why is the tail boom so long?", and "Why is the vertical fin so large?".  The first question is easier to answer than the second-- the tail boom itself creates relatively little drag, due to its slender cross-section and resulting low "wetted area".  To a first approximation, we can say that the tail boom is long so that the fin and horizontal tail don't need to be larger, to create the desired tail volume and flight characteristics.  In truth, the story is a little more complicated-- when we dive into the details of spiral stability theory, we find that extending the tail boom while keeping the "tail volume" the same actually does produce significant changes in the aircraft's flight characteristics, due to that fact that the airflow and "relative wind" are actually curved whenever the flight path is curved.1  This also means that for a given "tail volume", the drag reduction resulting from extending the tail boom and decreasing the tail area will be less in turning flight than in linear flight.
One reason that modern sailplanes are generally designed with much larger vertical tails than vintage sailplanes--or most single-engine general aviation aircraft-- is that the long slender tail booms used in modern sailplane designs have a very small cross-sectional area and thus contribute very little to the aircraft's yaw stability ("directional stability").  Nearly all of the required "weathervane effect" has to come from the tail itself.
Certainly the purpose of the long tail boom and large vertical fin is not to "balance" the weight of the cockpit.  Rather, the cockpit is placed in the optimum position to balance everything else.
Footnotes--

Re "the airflow and 'relative wind' are actually curved whenever the flight path is curved"-- another way to say this is to note that in curving flight, the aircraft is rotating as well as translating linearly, so at any given instant in time, different parts of the aircraft are moving through the airmass in different directions.  The instantaneous linear velocity at the nose and tail are different, and "tufts" or "telltales" at the nose and tail would not blow in the same direction.  Details of the effects of the curving relative wind are beyond the scope of this answer, but it may be interesting to compare and contrast the designs of competition-level radio-controlled model sailplanes versus "free-flight" model sailplanes.  Both have relatively long fuselages, but the former invariably have much larger vertical tails than the latter.  The latter must be designed to fly in stable circles with no pilot input, while the former are designed for good response to pilot control inputs.  The effects of the curving airflow in turning flight play a role in these dynamics.

Isn't there also a speed factor involved?  That is, the force on the tail is proportional to the airspeed, so a plane designed to fly slowly would need a larger tail to supply a given force,

One more tidbit of info: I was told that sailplanes have T-tails in order to be able to survive landing in fields with high grass or eg wheat.

Gliders spend a lot of time circling, entering thermals dodging others while ridge flying and generally performing manoeuvres designed to terrify most power pilots. Henry Ford remarked 'there is no substitute for litres'. For a glider pilot 'there is no substitute for span'. Mine is a titchy 15 meters. And even that represents a very large moment of inertia in the yaw axis. To enter circle you have to simultaneously increase the rate of yaw and pitch. Back on the pole then and a boot on the rudder (hopefully coordinated otherwise we go skidding all over the sky!). You can only get the moments needed from elevator (pitch) and rudder (yaw). The wing tips might offer a longer arm but there are issues with sticking control surfaces there. (Spin, anyone?)


text:
How does "continuous curvature" give stealth?
Wikipedia says

RCS reduction as a result of shape had already been observed on the Royal Air Force's Avro Vulcan strategic bomber,[117] and the USAF's F-117 Nighthawk. The F-117 used flat surfaces (faceting technique) for controlling radar returns as during its development (see Lockheed Have Blue) in the early 1970s, technology only allowed for the simulation of radar reflections on simple, flat surfaces; computing advances in the 1980s made it possible to simulate radar returns on more complex curved surfaces.[118] The B-2 is composed of many curved and rounded surfaces across its exposed airframe to deflect radar beams. This technique, known as continuous curvature, was made possible by advances in computational fluid dynamics, and first tested on the Northrop Tacit Blue.[119][114]

That makes no sense. The whole point of stealth shaping is to reduce the number of directions radar waves go. Faceting is the only way to achieve this.

The physics of radar detection are very complex, due to operating wavelengths being close to feature dimensions. Even in a faceted aircraft, radar waves are still reflected in every direction. The idea in all cases is to produce a distribution unfavorable for enemy radar.

"Faceting is the only way to achieve this." Woops, the stealth bomber folks really messed up. Please drop them a line and give them the bad news.

Electromagnetic radiation is reflected differently depending on wavelength and angle. While light will always be reflected back or scattered, with centimeter waves this only happens if the incidence is almost vertical. Only then will the radar pulse be thrown back like a tennis ball.
In the case of grazing incidence, the radio wave continues to travel through the material and is scattered at the edges. That is why edge alignment is so important for stealth. And this also determines the inclination of all surfaces: You will not find any vertical surface on a well-designed stealth airplane.
For the early modern designs (Have Blue, F-117, MBB Lampyridae) the computational capabilities (both computer power and the available codes!) only allowed to simulate plane surfaces. The result was the multi-faceted design with edge alignment even for the serrated edges of maintenance panels and landing gear doors. Especially the early Lockheed designs sported many edges and, consequently, their stealthiness was rather poor and only remedied by the generous application of RAM (radar absorbing material). With better computational methods, also curved surfaces could be simulated and it became possible to reduce edge count. But ideally the surface should be horizontal near all edges, which explains why fuselages and wings look like they had ironing edges. Note the pillow shape of the lateral edges of flaps on modern stealth airplanes: The same applies here, too.

Faceting is the only way to achieve this.

Not really. This only applied to the early designs of the Seventies and Eighties, and even then RAM helped a lot already. Generally, the number of edges should be reduced as much as practical and vertical surfaces avoided completely. While inclination against the horizontal is possible in the center of continuous surfaces, the surface orientation should be as horizontal as possible at all edges and the transition from the horizontal edge to the half-vertical center should be gradual.

Why is it more important for the surfaces to be horizontsl.near the edges than elsewhere?

@Abdullah So the incoming radiation hits the edge at a grazing angle and travels along the surface until it hits the next edge. Remember, all antennas of radar systems are positioned around the horizon if seen from the stealth aircraft.

@Abdullah If you don't mind that radiation is reflected back, the surfaces can be near-vertical. An inclination of 70° towards the horizon will reflect horizontal waves back at an angle of 40°, so very little of its energy will travel back to the receiver. But then you better don't roll that airplane to around 20° of bank angle.

Stealth capabilities of military aircraft (or any vehicles or apparatus for that matter) are not reliant on the form of the object only. Other major contributing factors are materials and joinery. Different materials can be used to absorb radar frequency radiation and joinery of parts is designed such that no favorable condition to for radiation to be reflected back towards its source is created (as in prisms for example.
The main goal of the form of a stealth object is to create a controlled RCS pattern that is directionally optimized such that for certain strategic directions the radar signature of the object is minimal. In other directions radar absorbing materials can be used to minimize the return pulse of the radar signal.
The following study simulates the radar signature of the F-22 Raptor fighter. Please note that this study does not represent the actual radar signature of the F-22 but it can be considered a close enough approximation of that
https://www.hindawi.com/journals/ijap/2018/5435837/
From the diagram you can see that the radar signature has substantial directional variation, so the F-22 is by no means invisible to radar at certain directions. However from directions that are considered necessary for successful missions completion the F22 is very, very difficult to see with radar.
The same applies for faceted designs such as the F-117. The difference between faceted and constant curve designs is that faceted designs have relatively large flat surfaces at certain directions and this provides strong radar return, whereas the curve designs give a return across larger sweep angle but the return is smaller.
For aircraft delivering ordinance it must be taken into account that releasing bombs, missiles and such will most certainly give out the position for the aircraft. Therefore if it for example turns away right after release, the radar signature during that maneuver is not that important, approach and departure signatures are. For loitering surveillance aircraft the design will be different.
Since modern materials and designs can produce low enough radar signatures for necessary directions using curved designs, they are preferred since they are by far superior aerodynamically. The preference for flight performance was actually the case already in the 60's, when the CIA overhead recommaisance program was commissioning a supersonic spy plane.Locheed A-12 was chosen over Convair Kingfish despite having inferior radar properties, because its flight performance was projected to be better. What A-12 and its successor SR-71 lacked in stealth, they made up in pure grunt. They simply were too fast for the missiles of the day.

Well, since you know "the only way to achieve" stealth is through faceting, then obviously the B-2, F-22, F-35, B-21, and quite a few drones must not be stealthy.
The Wikipedia article is weak.  They usually are for deeply technical articles.  Don't go there to learn real engineering or real depth.  Go there to learn a layperson's perspective on topics.
One of the problems with cutting edge military technologies is that people who know about it -- aren't allowed to talk about it.
Given that constraint, it seems unreasonable for Wikipedia to be a strong source of information.


text:
What is difference between TAA and MSA?
I know the 'TAA' usually used in RNAV chart, but sometimes it can be used in ILS chart as well. (When defining an IF, GPS is solely methods)
but While I'm studying, I can't find exact difference between TAA and MSA.
Can you let me know ?
BEST

From the FAA Chart User's Guide

Minimum Safe Altitude (MSA)
MSAs are published for emergency use on IAP charts. MSAs appear in the
  planview of all IAPs except on approaches for which a Terminal Arrival
  Area (TAA) is used. The MSA is based on the primary NAVAID, waypoint,
  or airport reference point on which the IAP is predicated. The MSA
  depiction on the approach chart contains the identifier of the
  NAVAID/waypoint/ airport used to determine the MSA altitudes. MSAs are
  expressed in feet above mean sea level and normally have a 25 NM
  radius; however, this radius may be expanded to 30 NM if necessary to
  encompass the airport landing surfaces.  Ideally, a single sector
  altitude is established and depicted on the planview of approach
  charts; however, when necessary to obtain relief from obstructions,
  the area may be further sectored and as many as four MSAs established.
  When established, sectors may be no less than 90° in spread. MSAs
  provide 1,000 feet clearance over all obstructions but do not
  necessarily assure acceptable navigation signal coverage.

Terminal Arrival Areas (TAAs)
The TAA icons will be positioned in the planview relative to their
  relationship to the procedure. The icon will not have feeder routes,
  airways, or radar vectors depicted. The TAA provides a transition from
  the enroute structure to the terminal environment with little required
  pilot/air traffic control interface for aircraft equipped with Area
  Navigation (RNAV) systems.  A standard TAA has three areas:
  straight-in, left base, and right base. The arc boundaries of the
  three areas of the TAA are published portions of the approach. A TAA
  provides minimum altitudes with standard obstacle clearance when
  operating within the TAA boundaries. TAAs are primarily used on RNAV
  approaches but may be used on an ILS approach when RNAV is the sole
  means for navigation to the IF; however, they are not normally used in
  areas of heavy concentration of air traffic.

Non-standard TAAs may also be published; i.e., one base leg, no base
  legs.

To summarize:
MSA's are referenced to a primary navaid, waypoint or airport reference point associated with the IAP. They show safe altitudes for a 360 degree area around the reference.
TAA's are referenced relative to their relationship to the procedure.  Generally, that means one reference diagram for each IAF placed near the IAF depiction.


text:
Please provide some expert commentary to the aborted flight of the German chancellor
This week the German chancellor's plane (an Airbus A340-300) had to abort the flight due to loss of communications. The newspaper write they only had a satellite phone left1:

Die Kommunikationssysteme waren ausgefallen, sodass kein Kontakt zum Boden möglich war. [Der Pilot] musste deshalb per Satellitentelefon mit dem Flughafen Köln/Bonn die ungeplante Landung vorbereiten.
The communication systems failed, contacting the ground wasn't possible.  The satellite phone had to be used to prepare the landing.
Man gehe derzeit von einem Fehler in einer elektronischen Verteilerbox aus, die sowohl die Funkanlage als auch das System zum Ablassen des Kerosins steuere, hieß es von der Luftwaffe.
Currently an error in a distribution box is most likely.  The box services the radio as well as the fuel dumping system.

Is the antenna for the satellite uplink an integral part of the air plane?  As suggested here for WiFi.  I assume satellite phone doesn't work from inside the plane without an external antenna, is this correct?
Maybe this is nitpicking, but when the satellite phone is build into the airplane, with the necessary tracking, why is this not considered part of the communications system?
How come that seemingly crucial yet quite simple two way radio can fail and is not powered by some very basic passive turbine?  How much power is approximately required by the radio?

I don’t know whether that’s relevant to that particular flight, but I expect a head-of-state aircraft to have many more communications systems than the ones normally on board a standard A340. It doesn’t say in the article whether they used an aeronautical or a „normal“ satellite phone (which would normally not be used by crew but by Ms Merkel and her staff). I somehow expect the latter, actually...?

Thanks, I didn't even know that there are two classes of satellite phones.  The articles are rather sparse with information.

Not sure this "question" is on topic. Because it is actually not a question...

How come this incident does not appear on avherald? https://avherald.com/

Article from Aviation Today with a little more detail.

@kebs Would you like the subject better, if I reformulate the it like this: What kind of satellite link was used during the following emergency and why was it an emergency?

Süddeutsche Zeitung has a side story today about a rodent on board which was seen by crew members back in october. The plane was parked on an island near to Bali then (earthquake danger?). They didn‘t captured it, according to the article. Do you remember „The fifth element“ ^^

A cellphone works well in an aircraft on the ground. It does not in the air due to distance and because cellphone network antennas don't look into the sky. But GPS of a cellphone works well during flight, but only at a window seat. I guess a satellite phone would work near a window, too.

@curious_cat, because AvHerald focuses on commercial aviation; it does occasionally cover accidents of other flights, but basically never just incidents.

Satellite phones would be for passenger use.  Over land, ATC uses VHF/UHF radio, and over sea, ATC uses HF radio or (text-based) satellite data links.
Based solely on the quotes in the question, it sounds like a power distribution panel in the aircraft failed, leaving them with no radios.  There are procedures for aircraft with failed radios to continue flight, but it's a lot of extra work for both the flight crew and ATC.  And who knows what other equipment on the same power panel was affected.
Having a satellite phone available means the crew may be able to contact ATC and arrange a diversion and landing faster and more easily than the usual no-radio (NORDO) procedures allow, but neither ATC nor the crew would want to operate that way for any longer than necessary to land safely.
Also, specific to a Head of State, the protection detail would want to reassure their comrades on the ground that the aircraft hasn't been attacked or hijacked, and they would also prefer any problems be repaired at home by people they've already vetted to reduce the risk of espionage or sabotage.